1. Field of the Invention
The invention is generally related to rockets and more particularly to solar thermal rockets.
2. General Background
Solar thermal rockets were first proposed in 1954 as a way to provide greater specific impulse than chemical rockets. These devices use the sun""s energy to heat a propellant (typically hydrogen) to extremely high temperatures and then expel the hot gas through a nozzle to provide thrust. The high temperature and low mass of the propellant combine to produce a specific impulse of two to four times that of a chemical rocket. Previously, solar thermal rockets have been of either a xe2x80x9cdirect gainxe2x80x9d design in which the propellant is heated directly by very large solar collectors during a long continuous burn, or of a xe2x80x9cthermal energy storagexe2x80x9d design which collects and stores energy from smaller collectors for use in short pulse burns. Each of these designs has its own advantages, but both have significant drawbacks that have prevented them from achieving commercial production.
The direct gain rocket uses a very large solar collector (concentrator) to heat hydrogen propellant as it passes through a cavity comprised of refractory metal tubes (typically rhenium). The advantage of this type of rocket is that the temperature of the propellant can be extremely high (theoretically greater than 3,000 degrees Kelvin), thus providing high specific impulse thrust. The problem with this design is that the solar collector must be extremely large (often one hundred to one hundred fifty feet in diameter) to provide enough energy to heat the propellant from its stored temperature of 300 degrees Kelvin to the desired temperature of greater than 3,000 degrees Kelvin. Concentrator technology has not matured to the point where such concentrators are available for space applications (i.e., light enough and in a small enough package to fit existing space launch vehicles) and it is arguable that this technology is still decades away.
The thermal energy storage design solves the concentrator problem by collecting and storing solar energy over an orbital period then using the stored energy to provide thrust for a short pulse burn. A number of these pulses are required to get the spacecraft to its destination. The longer the storage phase of the mission, the smaller the collector can be. This approach can enable the use of existing collector technology to develop a rocket. However, the drawback to such a system is that the energy storage materials (typically rhenium coated graphite or tungsten encapsulated boron nitride) have temperature limitations well below that of a direct gain rocket. Current designs are limited to about 2,400 degrees Kelvin, so the performance is well below that of a direct gain system. Solar rockets at 2,400 degrees Kelvin do not provide great enough performance margins over conventional chemical rockets to justify their development costs.
Thus, since the idea of a solar thermal rocket was first proposed, an operational system has yet to be developed. To make a practical system, one must find a way to reduce the size of the solar collectors without limiting the temperature of the receiver cavity.
The invention addresses the above need. What is provided is a solar thermal rocket that includes a solar energy receiver having two sections (a thermal energy storage section and a direct gain section), a solar concentrator, and a propulsion nozzle. In one embodiment, the focus of the solar energy between the storage section and the direct gain section is controlled by mechanical means such as movable insulation. In another embodiment, the focus of the solar energy between the storage section and the direct gain section is controlled by an optical switch in the form of relative motion between the solar concentrator and the solar energy receiver. Propellant is first heated by the thermal energy storage section and then the direct gain section before being directed to a propulsion nozzle.